The present invention relates generally to gas turbine engines, and, more specifically, to endurance testing of rotor blades therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel and ignited in a combustor for generating hot combustion gases. The gases flow downstream through turbine stages which extract energy therefrom for powering the compressor and producing useful work, such as powering a fan for propelling an aircraft in flight, for example.
In the compressor and turbines, stationary stator vanes direct air or gas between corresponding rotor blades for compressing the air or extracting energy from the combustion gases. Each rotor blade has an airfoil with an integral dovetail configured for being radially retained in a corresponding dovetail slot in the perimeter of a rotor disk. The dovetails may either be axial-entry or circumferential-entry, and are mounted in corresponding axial or circumferential dovetail slots in the perimeter of the corresponding rotor disks.
The blade dovetails may have various configurations due to the different blade size and function of the compressor and turbines. A typical dovetail includes a pair of lobes extending radially inwardly from a shank or neck of minimum cross section. The corresponding rotor disk includes a complementary dovetail slot defined between a pair of radially outwardly extending tangs. The dovetail lobes include radially outwardly facing contact or pressure faces which engage corresponding radially inwardly facing contact or pressure faces of the tangs.
During rotation of the blades in operation, centrifugal force or load is developed in each blade and is carried radially inwardly through the dovetails and reacted by the rotor disk. The centrifugal loads effect a nominal crush load over the corresponding pressure faces which causes a crush stress represented by the load divided by the contacting surface area of the pressure faces.
The rotor blades are also subject to aerodynamic loading by the air or combustion gases which flow over the corresponding airfoils during operation. For example, a multi-stage axial compressor includes respective rows of compressor blades decreasing in size for increasing pressure of the air in turn from stage to stage. Accordingly, the air pressure developed aft of the blade trailing edge is greater than the pressure forward of the blade leading edge, and effects a corresponding bending moment which is reacted by the dovetail lobes.
Accordingly, blade dovetails are designed in cooperation with their respective rotor disks for accommodating the various centrifugal and aerodynamic loads occurring during operation. Additionally, dovetail design must also accommodate thermal stresses due to the elevated temperatures experienced in the compressor as it pressurizes the air, as well as in the turbines heated by the hot combustion gases.
Compressor blades are typically designed for having an infinite life without undesirable cracking therein. For example, one type of large turbofan gas turbine engine has enjoyed many years of successful commercial use in this country, as well as abroad. The engine includes a high pressure compressor having several rows of circumferential entry compressor blades made of titanium.
The engine undergoes periodic field inspection of its various parts which has recently revealed undesirable cracking in some of the compressor blade dovetails having enjoyed high time operation with a substantial number of cycles and engine hours. The discovery of even a single crack requires the replacement of the entire row of compressor blades for ensuring optimum performance. Such cracking ends the useful life of the blades, which therefore fail to achieve the desired infinite life.
Accordingly, improved blade dovetail designs are being developed for addressing this high time dovetail cracking occurrence for further improving the engine design and reducing maintenance costs.
Modern three dimensional (3D) finite element analysis has been used to discover the source of high time compressor blade dovetail cracking. The dovetail neck, having minimum cross sectional area between the dovetail lobes, is initially the life limiting region of the dovetail having maximum local stress. The dovetail neck is therefore typically designed for suitably low stress to achieve infinite life.
However, analysis has uncovered that frictional shear forces develop during operation between the pressure faces and increase in magnitude as the blade accumulates cycles of operation. The friction force changes the loading profile experienced by the pressure faces and creates locally maximum contact stress at the two opposite edges of contact of each pressure face.
Accordingly the radially outer edge of contact located near the dovetail neck eventually experiences higher local stress than the neck itself in high time compressor blades. In particular, the forward lobe of a circumferential entry compressor blade dovetail experiences higher edge of contact stress than the aft lobe. And, the high-stress edges of contact provide crack initiation sites which may lead to premature dovetail failure.
The development of improved blade dovetails accommodating edge of contact stress necessarily requires suitable testing thereof for evaluating performance of the design. One conventional type of testing utilizes a whirligig in which an actual compressor rotor or spool is rotated to speed. The rotor includes an actual dovetail slot, such as a circumferential slot, in which is mounted an actual compressor blade for testing. Although the rotor may be operated at corresponding rotational speed to simulate actual operation in an engine, aerodynamic loading of the compressor blades is not included in the test. Such testing therefore has limited capability, and typically requires weeks or months of blade cycling to achieve sufficient high time cycles of operation.
Another conventional dovetail test includes an apparatus for pulling an individual compressor blade in tension in its corresponding dovetail slot. Tension loads are suitably cycled for achieving centrifugal loads and stress comparable to those obtained during actual rotary operation in an engine. However, the blade component pull test also fails to introduce aerodynamic loading. And, the dovetail slot is typically simulated using only a portion of an actual rotor disk which permits undesirable bending of the disk tangs leading to separation of the pressure faces in part near the outer edges of contact thusly changing the loading profile.
Although the whirligig testing and component pull testing may be effectively used for simulating dovetail neck stresses during operation, they are ineffective for simulating the complex loading profile at the pressure faces, and in particular at the edges of contact.
Accordingly, it is desired to provide an improved apparatus and method for more accurately simulating dovetail loading in cycle testing.